CENTRIFUGAL COMPRESSOR DIFFUSER PASSAGE BOUNDARY LAYER CONTROL

Abstract

A centrifugal compressor diffuser (42) includes a plurality of diffuser flow passages (22) extending through an annular diffuser housing (20) and circumferentially bounded by diffuser vanes (23) and axially bounded by forward and aft walls (101,100). A diffuser boundary layer bleed (96) for the passages may include boundary layer bleed apertures (106) or slots (130) disposed through the forward wall (101) and a downstream facing wall (142) canted at an acute cant angle to a downstream diffuser airflow direction (103) in the passages. Diffuser bleed flow (112) is bled from a diffuser boundary layer. Boundary layer bleed apertures can be located downstream of throat sections (28) of the flow passages near pressure sides of the vanes. A centrifugal compressor (18) may include the diffuser surrounding an annular centrifugal compressor impeller (32) and apparatus for flowing impeller bleed flow (102) from a radial clearance between an impeller tip (36) and a diffuser annular inlet (27) with diffuser bleed flow either mixed or separately to cool a turbine (16).

Description

BACKGROUND OF THE INVENTION

Technical Field

The present invention relates to bleed air from gas turbine engine centrifugal compressors.

One type of gas turbine engine includes a centrifugal compressor having a rotatable impeller to accelerate and, thereby, increase the kinetic energy of air flowing therethrough. A diffuser is generally located immediately downstream of and surrounding the impeller. The diffuser operates to decrease the velocity of the air flow leaving the impeller and transform the energy thereof to an increase in static pressure, thus, pressurizing the air.

A conventional gas turbine engine typically includes a compressor, combustor, and turbine, both rotating turbine components such as blades, disks and retainers, and stationary turbine components, such as vanes, shrouds, and frames routinely require cooling due to heating thereof by hot combustion gases. Cooling of the turbine, especially the rotating components, is important to the proper function and safe operation of the engine. It is known to bleed cooling air from the centrifugal compressor to help cool the turbine.

Failure to adequately cool a turbine disk and its blading, for example, by providing cooling air deficient in supply pressure, volumetric flow rate or temperature margin, may be detrimental to the life and mechanical integrity of the turbine. Depending on the nature and extent of the cooling deficiency, the impact on engine operation may range from relatively benign blade tip distress, resulting in a reduction in engine power and useable blade life, to a rupture of a turbine disk, resulting in an unscheduled engine shutdown.

Balanced with the need to adequately cool the turbine is the desire for higher levels of engine operating efficiency which translate into lower fuel consumption and lower operating costs. Since turbine cooling air is typically drawn from one or more stages of the compressor and channelled by various means, such as pipes, ducts, and internal passageways to the desired components, such air is not available to be mixed with fuel, ignited in the combustor and undergo work extraction in the primary gas flowpath of the turbine.

Total cooling flow bled from the compressor is a loss in the engine operating cycle and it is desirable to keep such losses to a minimum.

Some conventional engines employ clean air bleed systems to cool turbine components in gas turbines using an axi-centrifugal compressor as is done in the General Electric CFE738 engine. The turbine cooling supply air exits the centrifugal diffuser through a small gap between the diffuser exit and deswirler inner shroud. Other turbine cooling air methods include extracting cooling from the impeller or from a gap between the impeller and the diffuser exit.

U.S. Pat. No. 5,555,7211 to Bourneuf, et al, which issued on Sep. 17, 1996 and is entitled AGas Turbine Engine Cooling Supply Circuit@, discloses using bleed air from an impeller stage of a centrifugal compressor in a turbine cooling supply circuit for a gas turbine. U.S. Pat. No. 5,555,721 discloses impeller tip forward bleed flow and impeller tip aft bleed flow for cooling turbine components. U.S. Pat. No. 5,555,721 is assigned to the General Electric Co., the same assignee as this patent and is incorporated herein by reference.

U.S. Pat. No. 8,087,249 to Ottaviano, et al. which issued Jan. 3, 2012, and is entitled ATurbine Cooling Air From A Centrifugal Compressor@ discloses a gas turbine engine turbine cooling system including an impeller and a diffuser directly downstream of the impeller and a bleed for bleeding clean cooling air from downstream of the diffuser. U.S. Pat. No. 8,087,249 is assigned to the General Electric Co., the same assignee as this patent and is incorporated herein by reference.

Thus, there continues to be a demand for advancements in diffuser design and geometry that improves aerodynamic performance and reduces the overall engine radial envelope.

BRIEF DESCRIPTION OF THE INVENTION

A diffuser for a centrifugal compressor includes an annular diffuser housing, diffuser vanes axially extending between a forward wall and an aft wall of the diffuser housing, a plurality of diffuser flow passages extending through the housing and spaced about a circumference of the housing. The diffuser flow passages are bounded by the diffuser vanes and the forward and aft walls. A diffuser boundary layer bleed is provided for bleeding diffuser bleed flow from a diffuser boundary layer in each of the diffuser flow passages.

The diffuser boundary layer bleed may be configured for bleeding the diffuser bleed flow from the diffuser boundary layer at a position located in a region of flow weakness in each of the diffuser flow passages.

The diffuser boundary layer bleed may include boundary layer bleed apertures disposed through the forward wall. Each of the boundary layer bleed apertures may be a slot including a downstream facing wall angled or canted at an acute cant angle with respect to a downstream diffuser airflow direction in each of the diffuser flow passages respectively.

The boundary layer bleed apertures may be positioned or located downstream of throat sections of the diffuser flow passages near pressure sides of the diffuser vanes.

A centrifugal compressor including an annular centrifugal compressor impeller, a diffuser annularly surrounding the impeller, and a plurality of diffuser flow passages extending through a housing of the diffuser and spaced about a circumference of the housing. Each of the passages includes a throat section and a diffusing section downstream of the throat section. The diffuser flow passages are circumferentially bounded by diffuser vanes extending axially between forward and aft walls of the diffuser and a diffuser boundary layer bleed is provided for bleeding diffuser bleed flow from a diffuser boundary layer in each of the diffuser flow passages.

The centrifugal compressor may also include a radial clearance between an impeller tip of the impeller and an annular inlet of the diffuser, a means for mixing impeller bleed flow from the radial clearance with diffuser bleed flow from the boundary layer bleed apertures for providing turbine cooling air, and a means for flowing the turbine cooling air to a turbine or a means for flowing impeller bleed flow and the diffuser bleed flow separately to the turbine.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a sectional view illustration of a gas turbine engine with a centrifugal compressor for mixing impeller tip bleed flow and diffuser bleed flow in the compressor section before using the flows for cooling turbine components.

FIG. 2 is an enlarged sectional view illustration of the centrifugal compressor and a diffuser with diffuser bleed holes illustrated in FIG. 1.

FIG. 3 is an aft looking forward perspective view illustration of the diffuser and the diffuser bleed holes through 3-3 in FIG. 2.

FIG. 4 is an enlarged perspective view illustration of the bleed holes illustrated in FIG. 3.

FIG. 5 is a perspective view illustration of a portion of the diffuser and the diffuser bleed holes illustrated in FIG. 2.

FIG. 6 is an enlarged sectional view illustration of the centrifugal compressor tip and the diffuser bleed holes illustrated in FIG. 2.

FIG. 7 is a sectional view illustration of a gas turbine engine centrifugal compressor with an alternative arrangement for separately flowing impeller tip bleed for cooling turbine components.

FIG. 8 is a sectional view illustration of the gas turbine engine illustrated in FIG. 7 with an arrangement for separately flowing diffuser bleed flow for cooling turbine components.

FIG. 9 is an enlarged perspective view illustration of one of the impeller bleed flow ports illustrated in FIG. 7 and as taken through 9-9 in FIG. 10.

FIG. 10 is a forward looking aft perspective view illustration of an aft casing surrounding the centrifugal compressor and including the impeller and bleed flow ports illustrated in FIGS. 7 and 8 respectively.

FIG. 11 is cutaway perspective view illustration of impeller bleed flowpaths for one of the impeller bleed flow ports illustrated in FIGS. 7 and 9.

FIG. 12 is an enlarged perspective view illustration of one of the diffuser bleed flow ports illustrated in FIG. 8 and as taken through 12-12 in FIG. 10.

FIG. 13 is cutaway perspective view illustration of a diffuser bleed flowpath through one of the diffuser bleed flow ports illustrated in FIG. 8 and as taken through 12-12 in FIG. 10.

DETAILED DESCRIPTION OF THE INVENTION

Illustrated in FIG. 1 is a gas turbine engine high pressure centrifugal compressor 18 in a high pressure gas generator 10 of a gas turbine engine 8. The high pressure centrifugal compressor 18 is a final compressor stage of a high pressure compressor 4. The high pressure gas generator 10 has a high pressure rotor 12 including, in downstream serial or flow relationship, the high pressure compressor 14, a combustor 52, and a high pressure turbine 16. The rotor 12 is rotatably supported about an engine axis 25 by bearings in engine frames not illustrated herein.

The exemplary embodiment of the high pressure compressor 14 illustrated herein includes a five stage axial compressor 30 followed by the centrifugal compressor 18 having an annular centrifugal compressor impeller 32. Outlet guide vanes 40 are disposed between the five stage axial compressor 30 and the single stage centrifugal compressor 18. Compressor discharge pressure (CDP) air 76 exits the impeller 32 and passes through a diffuser 42 annularly surrounding the impeller 32 and then through a deswirl cascade 44 into a combustion chamber 45 within the combustor 52. The combustion chamber 45 is surrounded by annular radially outer and inner combustor casings 46, 47. Air 76 is conventionally mixed with fuel provided by a plurality of fuel nozzles 48 and ignited and combusted in an annular combustion zone 50 bounded by annular radially outer and inner combustion liners 72, 73.

The combustion produces hot combustion gases 54 which flow through the high pressure turbine 16 causing rotation of the high pressure rotor 12 and continue downstream for further work extraction in a low pressure turbine 78 and final exhaust as is conventionally known. In the exemplary embodiment depicted herein, the high pressure turbine 16 includes, in downstream serial flow relationship, first and second high pressure turbine stages 55, 56 having first and second stage disks 60, 62. A high pressure shaft 64 of the high pressure rotor 12 connects the high pressure turbine 16 in rotational driving engagement to the impeller 32. A first stage nozzle 66 is directly upstream of the first high pressure turbine stage 55 and a second stage nozzle 68 is directly upstream of the second high pressure turbine stage.

Referring to FIG. 1, the compressor discharge pressure (CDP) air 76 is discharged from the impeller 32 of the centrifugal compressor 18, used to combust fuel in the combustor 52, and to cool components of turbine 16 subjected to the hot combustion gases 54; such as, the first stage nozzle 66, first and second stage shrouds 71, 69 surrounding the first and second high pressure turbine stages 55, 56 respectively. The high pressure compressor 14 includes a compressor aft casing 110 and a diffuser forward casing 114 as more fully illustrated in FIGS. 1 and 2. The compressor aft casing 110 generally surrounds the axial compressor 30 and the diffuser forward casing 114 generally surrounds the centrifugal compressor 18 and supports the diffuser 42 directly downstream of the centrifugal compressor 18. The compressor discharge pressure (CDP) air 76 is discharged from the impeller 32 of the centrifugal compressor 18 directly into the diffuser 42.

Referring to FIGS. 2 and 3, the impeller 32 includes a plurality of centrifugal compressor blades 84 radially extending from a rotor disc portion 82. Opposite and axially forward of the compressor blades 84 is an annular blade tip shroud 90. The shroud 90 is adjacent to blade tips 86 of the compressor blades 84 defining a blade tip clearance 80 therebetween. The diffuser 42 includes an annular diffuser housing 20 having a plurality of tangentially disposed diffuser flow passages 22 extending radially therethrough, spaced about a circumference 26 of the housing 20, and through which diffuser airflow 103 flows in a downstream direction. Diffuser vanes 23 axially extend between a forward wall 101 and the aft wall 100 of the diffuser 42.

Referring to FIGS. 2 and 3, the diffuser vanes 23 circumferentially extend between adjacent ones of the diffuser flow passages 22. The diffuser flow passages 22 are partly defined and circumferentially bounded by the circumferentially spaced apart diffuser vanes 23. Adjacent ones of the passages 22 intersect with each other at radially inner inlet sections 24 of the passages 22 that define a quasi-vaneless annular inlet 27 of the diffuser 42. Each passage 22 further includes a throat section 28 downstream of and integral with the inner inlet section 24. Each passage 22 further includes a diffusing section 99 immediately downstream of the throat section 28.

Referring to FIGS. 2 and 6, a centrifugal compressor first cooling air source 92 for turbine cooling air 88 is a small predetermined radial clearance (C) located between an impeller tip 36 of the rotating impeller 32 and the annular inlet 27 of the static diffuser 42. Impeller bleed flow 102 from the radial clearance (C) is collected in a radially inner manifold 104. The predetermined. radial clearance (C) is designed to accommodate thermal and mechanical growth of the impeller 32 and is open to or in fluid communication with the radially inner manifold 104.

Referring to FIGS. 3-6, we have found that the diffuser airflow 103 on one side of the passage (such as passage 22) in multi-passage diffusers (such as the diffuser 42) that follow or are downstream of centrifugal impellers (such as the impeller 32) is often weak and may be subject to separation. Separation in the passage can generate high losses that lowers engine specific fuel consumption (SFC). This area or region of weak flow 127 is also believed to be a contributor to surge that limits flow range of the compressor.

A centrifugal compressor stage second cooling air source 94 for turbine cooling air 88 includes a diffuser boundary layer bleed 96 for bleeding diffuser bleed flow 112 from a diffuser boundary layer 113 in each of the diffuser flow passages 22 of the diffuser 42, illustrated herein as plurality of boundary layer bleed apertures 106. The diffuser boundary layer bleed 96, also referred to as fluidic bleed, helps reduce the weak flow and limit or prevent the unwanted flow separation. The diffuser boundary layer bleed 96 bleeds diffuser bleed flow 112 from the diffuser boundary layer 113 into a radially outer manifold 116.

The radially inner and outer manifolds 104, 116 are in fluid communication such that the impeller bleed flow 102 from the radially inner manifold 104 flows into the radially outer manifold 116. The impeller and diffuser bleed flows 102, 112 are mixed in the radially outer manifold 116 to provide the turbine cooling air 88 which is then ported or otherwise flowed from radially outer manifold 116 through a plurality of circumferentially distributed manifold ports 117 to the high pressure turbine 16. The turbine cooling air 88 may be channelled or flowed therefrom by external piping (not shown) to cool the first and second stage shrouds 71, 69 (illustrated in FIG. 1).

Substantially axially extending beams or struts 122 separate the radially inner and outer manifolds 104, 116 and the impeller bleed flow 102 passes between the struts 122 as it flows from the radially inner manifold 104 into the radially outer manifold 116. The fluidic bleed flow illustrated herein as the diffuser boundary layer bleed 96 represents a small amount of flow, less than 1% of the engine core flow. The fluidic bleed is strategically removed near the inception of the weak flow to improve the overall performance of the diffuser.

Referring to FIGS. 3-5, the boundary layer bleed apertures 106 may be holes or slots 130 through the forward wall 101 of the diffuser 42 as illustrated herein. The boundary layer bleed apertures 106 or slots 130 lead into and are in flow communication with the radially outer manifold 116. The slot 130 is positioned or located downstream of the throat section 28 near a pressure side 126 of the diffuser vane 23 at a position where the flow would begin to show weakness or instability in a diffuser without the diffuser boundary layer bleed 96. This position is located in what is referred to as a region of flow weakness 127. A slot width W may be sized with manufacturing constraints such as a minimum tool size. A slot length L may be selected to enable up to 3% of the engine core flow to be used.

The slot 130 should ideally be angled such that the diffuser bleed flow 112 exits the slot perpendicular to a forward surface 105 of the forward wall 101 of the diffuser 42 in a radial plane 132 passing through the engine centerline or axis 25 as illustrated in FIG. 5. However, because of constraints such as the slot extending through or very near a bend 134 in the forward wall 101 of the diffuser 42 this angle may be different. The slot 130 has radially outer and inner walls 136, 138, as illustrated in FIG. 6, and upstream and downstream facing walls 140, 142, as illustrated in FIGS. 4 and 5 respectively, extending through the forward wall 101. The downstream facing wall 142 is designed to scoop boundary layer air 144 in the diffuser boundary layer 113 only. Thus, the downstream facing wall 142 is angled or canted at an acute cant angle B of less than 90 degrees with respect to the diffuser airflow 103 (parallel to the direction boundary layer air 144 in the downstream direction in the diffuser flow passages 22 of the diffuser 42. It appears that an acute cant angle B of 45 degrees is desirable. However, the acute cant angle B is limited by geometry and manufacturing constraints on the outside of the diffuser so that an acute cant angle, for example about 22.5 degrees, is more practical.

Illustrated in FIGS. 7-13 is a gas turbine engine with a centrifugal compressor similar to the one illustrated in FIGS. 1-3 but with an alternative arrangement or design for separately gathering and flowing the impeller tip bleed and diffuser bleed flow for cooling turbine components. The impeller bleed flow 102 front the radial clearance (C), illustrated in FIG. 9, is flowed into and collected in a radially inner annular manifold 154 illustrated in FIGS. 7 and 9. Inter-manifold apertures 160 are disposed between the inner annular manifold 154 and a plurality of radially outer annular manifolds 156 illustrated in FIGS. 7, 9, and 13. The inter-manifold apertures 160 allow the impeller bleed flow 102 to flow front the inner annular manifold 154 into the outer annular manifolds 156. The impeller bleed flow 102 from the outer annular manifolds 156 is then ported or otherwise flowed through a plurality of circumferentially distributed impeller bleed flow manifold ports 157, illustrated in FIG. 10, to the high pressure turbine 16 for turbine cooling.

Referring to FIGS. 8, 10, and 11-13, the diffuser boundary layer bleed 96 bleeds diffuser bleed flow 112 from the diffuser boundary layer 113 into an annular diffuser bleed manifold 158 from where the diffuser bleed flow 112 is then ported or otherwise flowed through a plurality of circumferentially distributed diffuser bleed manifold ports 159 to the high pressure turbine 16 for turbine cooling. FIG. 10 illustrates the relative circumferential and axial locations of the impeller bleed flow manifold ports 157 and the diffuser bleed manifold ports 159 on and through the diffuser forward casing 114.

While there have been described herein what are considered to be preferred and exemplary embodiments of the present invention, other modifications of the invention shall be apparent to those skilled in the art from the teachings herein and, it is therefore, desired to be secured in the appended claims all such modifications as fall within the true spirit and scope of the invention. Accordingly, what is desired to be secured by Letters Patent of the United States is the invention as defined and differentiated in the following claims.

Claims

1. A gas turbine engine centrifugal compressor diffuser comprising: an annular diffuser housing,diffuser vanes axially extending between a forward wall and an aft wall of he diffuser housing,a plurality of diffuser flow passages extending through the housing and spaced about a circumference of the housing,the diffuser flow passages bounded by the diffuser vanes and the forward and aft walls, anda diffuser boundary layer bleed for bleeding diffuser bleed flow from a diffuser boundary layer in each of the diffuser flow passages.

2. The diffuser according to claim 1 further comprising the diffuser boundary layer bleed configured for bleeding the diffuser bleed flow from the diffuser boundary layer at a position located in a region of flow weakness in each of the diffuser flow passages.

3. The diffuser according to claim 1 further comprising the diffuser boundary layer bleed including boundary layer bleed apertures disposed through the forward wall.

4. The diffuser according to claim 3 further comprising each of the boundary layer bleed apertures being a slot including a downstream facing wall angled or canted at an acute cant angle with respect to a downstream diffuser airflow direction in each of the diffuser flow passages respectively.

5. The diffuser according to claim 3 further comprising the boundary layer bleed apertures positioned or located downstream of throat sections of the diffuser flow passages near pressure sides of the diffuser vanes.

6. The diffuser according to claim 5 further comprising each of the boundary layer bleed apertures being a slot including a downstream facing wall angled or canted at an acute cant angle with respect to a downstream diffuser airflow direction in each of the diffuser flow passages respectively.

7. A gas turbine engine centrifugal compressor comprising: an annular centrifugal compressor impeller,a diffuser annularly surrounding the impeller,a plurality of diffuser flow passages extending through a housing of the diffuser and spaced about a circumference of the housing,each of the passages including a throat section and a diffusing section downstream of the throat section,the diffuser flow passages circumferentially bounded by diffuser vanes extending axially between forward and aft walls of the diffuser, anda diffuser boundary layer bleed for bleeding diffuser bleed flow from a diffuser boundary layer in each of the diffuser flow passages.

8. The centrifugal compressor according to claim 7 further comprising the diffuser boundary layer bleed configured for bleeding the diffuser bleed flow from the diffuser boundary layer at a position located in a region of flow weakness in each of the diffuser flow passages.

9. The diffuser according to claim 7 further comprising the diffuser boundary layer bleed including boundary layer bleed apertures disposed through the forward wall.

10. The centrifugal compressor according to claim 9 further comprising each of the boundary layer bleed apertures being a slot including a downstream facing wall angled or canted at an acute cant angle with respect to a downstream diffuser airflow direction in each of the diffuser flow passages respectively.

11. The centrifugal compressor according to claim 10 further comprising the boundary layer bleed apertures positioned or located downstream of throat sections of the diffuser flow passages near pressure sides of the diffuser vanes.

12. The centrifugal compressor according to claim 11 further comprising each of the boundary layer bleed apertures being a slot including a downstream facing wall angled or canted at an acute cant angle with respect to a downstream diffuser airflow direction in each of the diffuser flow passages respectively.

13. The centrifugal compressor according to claim 9 further comprising: a radial clearance between an impeller tip of the impeller and an annular inlet of the diffuser,a means for mixing impeller bleed flow from the radial clearance with the diffuser bleed flow from the boundary layer bleed apertures for providing turbine cooling air and flowing the turbine cooling air to a turbine, ora means for flowing the impeller bleed flow and the diffuser bleed flow separately to the turbine.

14. The centrifugal compressor according to claim 13 further comprising each of the boundary layer bleed apertures being a slot including a downstream facing wall angled or canted at an acute cant angle with respect to a downstream diffuser airflow direction in each of the diffuser flow passages respectively.

15. The centrifugal compressor according to claim 13 further comprising the boundary layer bleed apertures positioned or located downstream of throat sections of the diffuser flow passages near pressure sides of the diffuser vanes.

16. The centrifugal compressor according to claim 15 further comprising each of the boundary layer bleed apertures being a slot including a downstream facing wall angled or canted at an acute cant angle with respect to a downstream diffuser airflow direction in each of the diffuser flow passages respectively.

17. The centrifugal compressor according to claim 9 further comprising: a radial clearance between an impeller tip of the impeller and an annular inlet of the diffuser,the radial clearance in fluid communication with a radially inner manifold,the boundary layer bleed apertures in flow communication with a radially outer manifold,the radially inner manifold in fluid communication with the radially outer manifold such that the impeller bleed flow flows into the radially outer manifold and mixes with the diffuser bleed flow to form turbine cooling air, andmeans for flowing turbine cooling air out of the radially outer manifold.

18. The centrifugal compressor according to claim 17 further comprising each of the boundary layer bleed apertures being a slot including a downstream facing wall angled or canted at an acute cant angle with respect to a downstream diffuser airflow direction in each of the diffuser flow passages respectively.

19. The centrifugal compressor according to claim 18 further comprising the boundary layer bleed apertures positioned or located downstream of throat sections of the diffuser flow passages near pressure sides of the diffuser vanes.

20. The centrifugal compressor according to claim 19 further comprising each of the boundary layer bleed apertures being a slat including a downstream facing wall angled or canted at an acute cant angle with respect to a downstream diffuser airflow direction in each of the diffuser flow passages respectively.

21. The centrifugal compressor according to claim 9 further comprising: a radial clearance between an impeller tip of the impeller and an annular inlet of the diffuser,the radial clearance in fluid communication with a radially inner annular manifold,inter-manifold apertures disposed between the inner annular manifold and a plurality of radially outer annular manifolds,a means for porting and flowing the impeller bleed flow from the radial clearance through a plurality of circumferentially distributed impeller bleed flow manifold ports in and through an diffuser forward casing surrounding the centrifugal compressor to the high pressure turbine for turbine cooling,the diffuser boundary layer bleed in fluid flow communication with and operable for bleeding the diffuser bleed flow into an annular diffuser bleed manifold, anda means for porting and flowing the diffuser bleed flow through a plurality of circumferentially distributed diffuser bleed manifold ports in and through the diffuser forward casing to the high pressure turbine for turbine cooling.

22. The centrifugal compressor according to claim 21 further comprising each of the boundary layer bleed apertures being a slot including a downstream facing wall angled or canted at an acute cant angle with respect to a downstream diffuser airflow direction in each of the diffuser flow passages respectively.

23. The centrifugal compressor according to claim 22 further comprising the boundary layer bleed apertures positioned or located downstream of throat sections of the diffuser flow passages near pressure sides of the diffuser vanes.

24. The centrifugal compressor according to claim 23 further comprising each of the boundary layer bleed apertures being a slot including a downstream facing wall angled or canted at an acute cant angle with respect to a downstream diffuser airflow direction in each of the diffuser flow passages respectively.